The present invention concerns a flight control device for an aircraft, in particular helicopter.
The flight control device of an aircraft captures command instructions generated by the pilot and transmits it to the aerodynamic control surfaces that cause a change in the spatial orientation of the craft. In first generation aircraft the flight control device was mechanical, pilot commands being generated by joysticks and rudder bars connected to the control surfaces by mechanical linkages based on links and/or cables. Second generation aircraft had a system for amplifying the force applied by the pilot integrated into the control surface, in the form of an actuator. The actuator was operated by the linkage and its movement was transmitted to the control surface. As safety requirements were strengthened, it was necessary to duplicate the linkages to survive a link or a cable breaking.
In today's third generation aircraft, to which the present invention applies, mechanical linkages are replaced with electrical wires that transmit the position of control devices actuated by the pilot to computers that transform them for execution by the actuators of the appropriate control surfaces.
The immediate benefit of a solution of this kind, representing an electric flight control device for aircraft in particular and more particularly for civil aircraft, is a significant saving in weight and the ease of increasing the number of command instruction routing paths. Safety is enhanced because a plurality of electrical paths in different locations in the craft are less vulnerable than one or two mechanical paths. The control computers also generate command instructions for stabilizing the aircraft and these are superposed on the command instructions from the pilot.
Although the saving in weight obtained by the use of an electrical flight control device is less on a helicopter than on a fixed wing aircraft, the benefit of a device of this kind on a helicopter is far from negligible, in particular because of the multiplicity of paths that are possible.
Nevertheless, when applied to a helicopter, a flight control device must offer particularly high performance, in particular ensuring safety with one control surface per axis and by being reconfigurable immediately in the event of a failure. In this respect fixed wing aircraft and helicopters differ in two important aspects:
firstly, a fixed wing aircraft includes a plurality of control surfaces that can cause it to react about the same axis. Accordingly, if the control path for one control surface fails, the path for command instructions to the other control surface can take over. In contrast, the helicopter has only one control surface per axis and the control of this control surface must assure safety on its own; and PA1 on the other hand, a fixed wing aircraft has some degree of inherent stability, allowing the control surface control system to be down for some time period during reconfiguration after a failure. In contrast, a helicopter is inherently unstable, which imposes continuous control at the risk of destabilizing it. PA1 a plurality of n instruction generating systems each of which generates a set of first command instructions for said control units and at least one of which performs autosurveillance and generates corresponding surveillance signals; and PA1 a plurality of p servocontrol systems each of which: PA1 the second instructions determined by at least some of the p servocontrol systems being communicated to the control units of the aircraft as command instructions. PA1 of the analogue type; or PA1 of the digital type; or PA1 partly of the analogue type and partly of the digital type.
Some control modes are also very demanding in terms of computation capacity as they integrate, over and above command instructions from the pilot, input on the status of the helicopter, for example attitude or angular speed, in order to improve control performance, and in particular stability. The necessary computing power can then be provided only by digital computers. However, digital computers can fail for various reasons (hardware failure, software error, sensitivity to ionizing radiation).
One method of avoiding such failures is to use computers with different technologies for each path. In particular analogue technology is a good response to such problems. However, its use must be restricted because its low computing power can handle only relatively impoverished control modes and this is acceptable only if the probability of these occurring is low for mission reliability reasons.